Pneumatic resolver for missile control

ABSTRACT

A pneumatic attitude control system for an artillery-type missile which is spin stabilized during flight. The control system includes a sensor portion which is roll stabilized and, thus, will not roll with the missile. Thus, attitude angles measured by the roll stabilized sensors will be space fixed. Control valves for providing restoring torques on the missile are mounted on the rotating part of the missile for rotation therewith. A resolution system is provided for resolving the space-fixed signals from the roll-stabilized sensors into the rolling missile&#39;&#39;s coordinate system for desired operation of the control valves.

llnited States Patent [72] Inventor Charles Lynn Lewis Huntsville, Ala. [2]] Appl. No. 839,829 [22] Filed July 8, 1969 [45] Patented Oct. 19, 1971 [73] Assignee The United States of America as represented by the Secretary of the Army [54] PNEUMATIC RESOLVER FOR MISSILE CONTROL 9 Claims, 9 Drawing Figs.

[52] US. Cl 244/3.22, 137/815 [51] Int. Cl ..F42b 15/02, G06f l/OO, Fl5d 1/02 [50] Field of Search 244/322 [56] References Cited UNITED STATES PATENTS 3,278,140 10/1966 Evans 244/322 ATTITUDE ANGLE RESOLVER 8Y5.

ATTITUDE ANGLE RESOLVER SYS SPACE /XE: PZRTlON OF M SS LE 3,325,121 6/1967 Ba nszaketal. 3,502,285 3/1970 Gambill Primary Examiner- Benjamin A. Borchelt Assistant ExaminerThomas H. Webb Attorneys-Harry M. Saragovitz, Edward J. Kelly, Herbert Berl and Harold W. Hilton ROLLING PORTION 'JF MISS LE H PATENTEDBET 19 I97! SHEET u c s FIG. 5

Charies L. Lewis,

PATENTEBUCI 19 Ian 3,614,027 SHEET 5 OF 6 PICKOFF (29) PLATE ROTATION FIG. 8

l-IJ a";

x \VALUE OF P SW2 2" P2 FOR NO INPUT E: D O

PLATE ROTATION FIG. 7

Charms L. Lewis,

J INVENT R 22/ wt; BY L%/ PNEUMATIC RESOLVER FOR MISSILE CONTROL SUMMARY OF THE INVENTION A pneumatic control system for a missile having a portion thereof which is spun prior to launch and a nonspinning (rollstabilized) portion. Tubes are mounted on the roll-stabilized portion for carrying pneumatic signals therefrom representing space-fixed attitude angles. The signals are obtained from a pneumatic gyro and carried by the tubes to exhaust perpendicularly onto a cam or plate having concentric grooves, the depths of which vary in a sinusoidal manner around the circumference. The cam is mounted to a rotating portion of the missile. Thus, the distance between the ends of the tube and the surfaces of the plate varies as the plate rotates relative to the ends of the tubes. The variable depths of the grooves change the back pressure in each tube and thereby change the pressure in a pickoff line upstream of the exhaust point, or end of the tubes. The pickoff pressure is proportional to the distance between the cam surface and the tube and will change the pressures in the pickoff lines. The shape of the cam provides for the occurrence of pressure variations in a sinusoidal manner.

The outputs of the back pressure pickofis are summed with necessary bias signals to provide analog signals which provide the required resolution of the space-fixed attitude angles into rolling coordinates.

It is, therefore, an object of the present invention to provide a pneumatically controlled attitude control system for a missile.

It is another object of the present invention to provide a missile attitude control system having a sensor portion thereof which is space fixed and a control portion thereof which is rotating during missile flight.

It is a further object of this invention to provide such a control apparatus with means to resolve signals from the spacefixed sensor into the missiles rolling coordinate system for utilization by the control portion of the system.

These and other objects of the present invention will be more readily apparent from the following detailed description and accompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a diagrammatic representation of the control system of the present invention.

FIG. 2 is a diagrammatic representation of the control system including the attitude angle resolver system as shown in FIG. 1.

FIG. 3 is a sectional view along line 3-3 of FIG. 1 showing the resolver cam and slipring arrangement with the tubes entering the slipring assembly removed for clarity.

FIG. 4 is a diagrammatic view of the space-fixed and rolling body coordinate systems.

FIG. 5 is a diagrammatic view of mechanism for coordinate transformation of the space-fixed attitude angle in a first attitude plane.

FIG. 6 is a diagrammatic view of mechanism for coordinate transformation of the space-fixed attitude angle in a second attitude plane.

FIG. 7 is a graphical representation illustrating pickofi' pressure variation versus plate rotation.

FIG. 8 is a graphical representation illustrating output pressure versus plate rotation.

FIG. 9 is a pictorial view illustrating diagrammatically, the missiles roll-stabilized and rotating portions and the slipring connection therebetween.

DESCRIPTION OF THE PREFERRED EMBODIMENT As shown in FIG. 1, a missile 10 is shown to include a rollstabilized portion 12 and a rotating portion 14 which are secured together for relative rotation therebetween by means well known in the art. A control system 16 carried in the missile includes a rotating control portion 18 and a space-fixed sensor portion 20. Sensor portion 20 is shown to include a plurality of tubes 22-29 which are disposed in communication with a rotating cam 15 which is carried by, and disposed for rotation with rotating portion 14 of the missile. Cam 15 is provided with a pair of concentric surfaces 11 and 13 (see FIG. 3) of varying depth. Tubes 22-29 communicate between a pair of attitude angle resolver systems 31 and 33 and cam 15 for generation of signals which are resolved into the missiles rolling coordinate system by resolver systems 31 and 33 and transferred to a pulse duration modulation mechanism 17 and through a fluidic slipring 19 to a plurality of control valves 21 mounted on the periphery of the missile. Input signals to resolver systems 31 and 33 are received from a pneumatic gyro 9.

FIG. 2 is a block diagram of the control system of FIG. 1 and illustrates the mechanism of the attitude angle resolver systems as generally shown in FIG. 1. As shown in FIG. 2, the attitude angle resolver system includes a plurality of outputs 30-37 of a plurality of center dump amplifiers 40-47, respectively, which communicate, respectively, with tubes 48-55 are connected to tubes 22-29 intermediate the input center dump amplifiers 40-47 and rotating cam 15. Tubes 48, 49, 50 and 51 communicate respectively, with tubes 22, 23, 24 and 25 and a plurality of summing amplifiers 58, 59, and 60, respectively Lines 52, 53, 54, and 55 connect, respectively intermediate lines 26, 27, 28 and 29 and a second plurality of summing amplifiers 61, 62, and 63.

Amplifiers 41, 43, 45 and 47, are provided with a second plurality of output tubes (FIGS. 2, 5 and 6) 68-70, 72-74, 76-78, and 80-82, respectively, (FIGS. 5 and 6) which are provided with an attenuator or orifice 84 to reduce pressure in these tubes by three-fourths. This serves to provide the input sine wave with a polarity. The outputs from the bank of output summing amplifiers (FIG. 2) are transmitted through tubes 86, 88, and 92 to a pulse duration modulation mechanism 17 to produce a pulse duration modulation signal which is transferred through slipring" type mechanism 19 (FIG. 9). The transferred control signals are then used to drive the plurality of control valves 21 carried around the periphery of the rotating portion of the missile.

As more clearly seen in FIG. 5, the mechanism for transformation of space-fixed signals for control in one plane F0 is illustrated, assuming this to be the missles body-fixed pitch control plane. In this control mechanism, a first pair of input amplifiers 40 and 41 are connected to input sources and 102 which provide a control signal across the power jet flowing through entrance 104, which is received from a missile carried source, not shown, of amplifiers 40 and 41. The jet is disposed for flow through tubes 24 and 25 of amplifier 40 and through tubes 68 and 70 of amplifier 41. Tubes 24 and 25 terminate adjacent the cam surface 15. However, pickoffs 50 and 51 connect into tubes 24 and 25, respectively, upstream of the ends of the tubes which lie adjacent cam 15 and connect to the power jet entrance 104 of a summing amplifier 58. Tubes 68 and 70 of amplifier 41 connect adjacent power jet inlet 104 of amplifier 58, as shown in FIG. 5.

Outputs I06 and 108 of amplifier 58 connect across opposite sides of entrance 104 of another output summing amplifier 60 which is provided with output tubes 90 and 92 which connect to the pulse duration modulation mechanism 94 (FIGS. land 2).

In like manner, input amplifier 42, as shown in FIG. 5, is connected to inputs 110 and 112 which provide a control signal across the power jet flowing through entrance 104 of amplifiers 42 and 44. The jet is disposed for flow through tubes 22 and 23 of amplifier 42 and through'tubes 72 and 74 of amplifier 43. Tubes 22 and 23 terminate adjacent cam surface 15. However, pickoffs 48'and 49 connect into tubes 22 and 23 upstream of the ends of the tubes which lie adjacent cam 15 and these pickoffs connect adjacent the entrance 104 of a summing amplifier 59. Outputs 110 and 112 of summing amplifier 59 connect across entrance 104 of. output summing amplifier 60.

Similarly, to provide control in the plane rt), assuming this to be the missiles body-fixed yaw control plane attitude in yaw, the control mechanism as illustrated in FIG. 6 is utilized. In this control mechanism the pair of input center dump amplifiers 44 and 45 are connected to input sources 1 14 and 116 which provide a signal across the power jet flowing through the power jet inlet 104 of amplifiers 44 and 45. The jet is disposed for flow through tubes 26 and 27 of amplifier 44. Tubes 26 and 27 terminate adjacent cam surface 15. However, pickoffs 52 and 53 connect into tubes 26 and 27 upstream of the ends of the tubes which lie adjacent cam and these pickoffs connect adjacent the power jet inlet 104 of output summing amplifier 61. Outputs 118 and 120 of summing amplifier 61 connect across inlet 104 of an output summing amplifier 63 which is provided with a pair of output tubes 90 and 92 which connect into pulse duration modulation mechanism 17.

In like manner, input amplifier 46, as shown in FIG. 6, is connected to inputs 122 and 124 which provides a control signal across the power jet flowing through inlet 104 of amplifiers 46 and 47. The jet is disposed for flow through tubes 28 and 29 of amplifier 46 through tubes 80 and 82 of amplifier 47. Tubes 28 and 29 terminate adjacent cam surface 15. However, pickoffs 54 and 55 connect into tubes 28 and 29 upstream of the ends of the tubes which lie adjacent cam 15 and these pickofis connect adjacent the entrance 104 of output summing amplifier 62. Outputs 126 and 128 of output summing amplifier 62 connects across inlet 104 of output summing amplifier 63.

The resolution required in the control system of the present invention is illustrated in FIG. 4, where the solid axis represent the space-fixed coordinate system and the dotted axes represent the body-fixed coordinate system. The angle through which the missile with roll frequency to has rolled in time I is denoted mt.

Assuming 0 is the missiles space-fixed attitude angle in pitch and (l: is the space-fixed attitude angle in yaw, then the rolling (body-fixed) signals are seen to be sin-cos functions of these angles and are given by:

I8=0 cos wt-l-tll sin wt IilFtlr cos wt-O sin out.

To mechanize this pneumatic resolution, tubes which are mounted on the roll-stabilized portion of the control system and which carry analog (smoothly varying) pneumatic signals representing the space-fixed attitude angles 0 and ill are positioned so they exhaust perpendicularly into concentric grooves cut into a plate that is affixed to the rolling portion of the missile. This is shown in FIG. 2. The depth of the grooves cut into the plate varies in a sinusoidal manner around the circumference.

From fluid flow theory, it is known that the back pressure in the tube (which is picked off upstream of the exhaust point) is proportional to the distance x between the tube end and the bottom of the groove. The variable depth of the groove will change the pressure in the pickoff lines as follows. Assume that in FIG. 5, which shows the mechanization of one of the resolution equations, line 29 has a constant freeflow pressure of X p.s.i. When the line end is located at the 0 position of the plate, the pressure at pickoff 55 will also be X p.s.i. This indicates that the distance between groove bottom and tube end at the 0 location should be zero. As the plate rotates until the 90 position is under line 29, the pickoff pressure drops to 3X/4 p.s.i. Then, the pressure at the pickoff continues to drop as the plate rotates toward 180. At 180, the pressure at the pickoff is X2 p.s.i. The pressure then begins to build up through the next 180 of motion until it is again X p.s.i. after one complete cycle of the plate. This ratio of decrease and increase of the pressure from its original value during a cycle of the plate will hold for all of the lines impinging upon the plate. The depth of the grooves is to be out such that the above pressure variations will occur in a sinusoidal manner. This is shown in FIG. 7.

By locating the impingement points (points where tubes exhaust into the grooves) in a certain manner, as shown in FIGS. 5 and 6, and summing the outputs of the back pressure pickoffs with necessary bias signals, analog signals can be generated which are the required resolution of the space-fixed attitude angles 0 and ill into coordinates. The summation of the pickoff signals and the bias signals can be perfonned by any commercial pneumatic summing device.

One representative cycle of the pneumatic resolver would go as follows: First, it is assumed that with no input signal if =P21, P =P there will be no pressure in the lines that exhaust in to the grooves, and thus the output will correctly remain zero (P,=P regardless of the orientation of the plate. This is accomplished by the use of the center dump input amplifiers. For purposes of explaining the operation of the system, in FIG. 5, let P,=P =Y p.s.i. for no input. The actual magnitude is not critical to the operation of the resolver. Next, assume the input is a positive pitch attitude angle 6 which causes a differential pressure of P P a=+X p.s.i. across the center dump input amplifiers. In addition, assume the yaw input is zero (P P2,r=0). The sign of the input is given by the input amplifiers. The sign changes inherent in the sin-cos terms as the plate rotates in l, 2 quadrant Sin wr= in 3, 4 quadrant in l, 4 quadrant Cos m1= in 2, 3 quadrant is taken care of by the bias, which is set to the value of the pickoff pressure at the and 270 points. This value, as previously defined is 3X/4 p.s.i., where X is the input pressure.

The change in the output pressure P, and P of FIG. 5 as the plate rotates through one cycle is shown in FIG. 8. Examination of the F0 equation (with 6=positive quantity, vll=zero) indicates that F0=positive value (P, P for 0 to 90, zero at 90 (P,=P negative for 90 to 270 (P P,), zero at 270, and positive for 270 to 360 (P, P This is precisely what FIG. 8 shows the pneumatic resolver does. The exact magnitude of the output differential pressure as compared with the input differential pressure magnitude will depend upon the gain set in the pneumatic amplifiers and summers.

The function in the pneumatic resolver is to generate the F0 and Fill signals from 0 and ill signals. Once these signals are obtained, they may be summed with a high-frequency carrier signal to produce a pulse duration modulation control signal. This pneumatic signal, now in the form of a pulse train, is then transferred physically in a slipring" fashion from the roll-stabilized portion of the missile to the rolling portion via a fluidic switching slipring 19, as shown in FIG. 9.

The fluidic slipring is utilized to transfer the PDM control pulse train from the roll-stabilized portion to the rolling missile section of operation of the control valves. The axle shaft 129 of the roll-stabilized section is mounted in bearings 130 as shown in FIG. 9 to an opening 131 of the rolling portion of the missile. The axle shaft has a plurality of grooves 132 cut circumferentially therearound which are fed by the outputs 133, 134, 135, 136, ofthe PDM control mechanism 17. Tubes 137, 138, 139 and 140 mounted in the rolling portion of the missile are disposed for communication with the grooves and are disposed to pick up pressure in the grooves of the shaft, and transfer these pressures to control valves 21.

It is to be understood that the cam 15 may be a plate having a pair of grooves cut therein and provided with surfaces that vary sinusoidally, or, if desired, the cam may be a member wherein the cam surfaces protrude above a main surface in a manner to provide the sinusoidal output from the tubes. Additionally, the gyro for 9 referred to in this application may be any of the many commercially available type pneumatic gyros. One such type gyro, for example, may be obtained from General Precision, lnc., Singer-Kearfott Division, Little Falls, New Jersey.

The pulse duration modulation principle as utilized in the present invention is similar to that disclosed in the patent to Kenneth C. Evans, entitled "Pure Fluid Amplifier and Pure Fluid Amplifier Attitude Control System for Missiles No. 3,278,140, issued Oct. ll, 1966. Additionally, pulse duration modulation principle is disclosed by Hancock, in An Introduction to the Principle of Communication Theory", Mc Graw-Hill 1961.

lclaim:

1. An attitude control system for a missile having spinning and nonspinning portions, said attitude control system includmg:

a. sensing means carried in said nonspinning portion and disposed for sensing the attitude of said missile and generating space-fixed signals indicative thereof;

b. resolver means disposed for resolving said attitude signals into the coordinate system of said spinning portion;

c. control means disposed for receiving said resolved signals for utilization thereof to provide resorting torques on said missile.

2. An attitude control system as set forth in claim 1 including pulse duration modulation means connected between said resolver means and said control means to transform said signals into pulses for actuation of said control means.

3. An attitude control system as set forth in claim 2 wherein said sensing means includes a pneumatic gyro disposed for generating signals indicative of missile attitude.

4. An attitude control system as set forth in claim 3 wherein said resolver means includes:

a. a plurality of fluid amplifiers connected to said gyro to receive input signals therefrom; each of said amplifiers having a pair of output channels;

b. cam means carried on said spinning portion of said missile for coaction with the outputs of predetermined ones of said fluid amplifiers for generating back pressures therein, said pressures varying in response to rotation of said spinning portion.

5. An attitude control system as set forth in claim 4 wherein said cam means includes a pair of tracks having surfaces of varying depths, said surfaces varying in a sinusoidal manner.

6. An attitude control system as set forth in claim 5 wherein a first plurality of fluid amplifiers are disposed for providing signals for control of said missile in a first attitude plane, and a second plurality of fluid amplifiers are disposed for providing signals for attitude control in a second plane, said first plurality of fluid amplifiers having predetermined ones thereof provided with outputs disposed adjacent the first of said cam tracks, said second plurality of fluid amplifiers having predetermined ones thereof provided with outputs disposed adjacent the second of said cam tracks.

7. An attitude control system as set forth in claim 6 wherein said first and second plurality of amplifiers includes:

a. pickoff means connected to the output of said predetermined ones of said fluid amplifiers intermediate said cam and the body of said fluid amplifiers; and,

b. a plurality of fluid summing amplifiers having said pickoff means connected across the inlet thereof;

c. an output summing amplifier having the outputs of said summing amplifier connected across the inlet thereof.

8. An attitude control system as set forth in claim 7 wherein said pulse duration modulation means is connected to the outputs of said output summing amplifiers to receive signals therefrom.

9. An attitude control system as set forth in claim 8 including fluidic slipring means connected between said spinning and nonspinning portions of said missile to transfer signals from said pulse duration modulation means to said control means. 

1. An attitude control system for a missile having spinning and nonspinning portions, said attitude control system including: a. sensing means carried in said nonspinning portion and disposed for sensing the attitude of said missile and generating space-fixed signals indicative thereof; b. resolver means disposed for resolving said attitude signals into the coordinate system of said spinning portion; c. control means disposed for receiving said resolved signals for utilization thereof to provide resorting torques on said missile.
 2. An attitude control system as set forth in claim 1 including pulse duration modulation means connected between said resolver means and said control means to transform said signals into pulses for actuation of said control means.
 3. An attitude control system as set forth in claim 2 wherein said sensing means includes a pneumatic gyro disposed for generating signals indicative of missile attitude.
 4. An attitude control system as set forth in claim 3 wherein said resolver means includes: a. a plurality of fluid amplifiers connected to said gyro to receive input signals therefrom; each of said amplifiers having a pair of output channels; b. cam means carried on said spinning portion of said missile for coaction with the outputs of predetermined ones of said fluid amplifiers for generating back pressures therein, said pressures varying in response to rotation of said spinning portion.
 5. An attitude control system as set forth in claim 4 wherein said cam means includes a pair of tracks having surfaces of varying depths, said surfaces varying in a sinusoidal manner.
 6. An attitude control system as set forth in claim 5 wherein a first plurality of fluid amplifiers are disposed for providing signals for control of said missile in a first attitude plane, and a second plurality of fluid amplifiers are disposed for providing signals for attitude control in a second plane, said first plurality of fluid amplifiers having predetermined ones thereof provided with outputs disposed adjacent the first of said cam tracks, said second plurality of fluid amplifiers having predetermined ones thereof provided with outputs disposed adjacent the second of said cam tracks.
 7. An attitude control system as set forth in claim 6 wherein said first and second plurality of amplifiers includes: a. pickoff means connected to the output of said predetermined ones of said fluid amplifiers intermediate said cam and the body of said fluid amplifiers; and, b. a plurality of fluid summing amplifiers having said pickoff means connected across the inlet thereof; c. an output summing amplifier having the outputs of said summing amplifier connected across the inlet thereof.
 8. An attitude control system as set forth in claim 7 wherein said pulse duration modulation means is connected to the outputs of said output summing amplifiers to receive signals therefrom.
 9. An attitude control system as set forth in claim 8 including fluidic slipring means connected between said spinning and nonspinning portions of said missile to transfer signals from said pulse duration modulation means to said control means. 